The present invention relates generally to coolant supply systems in gas turbine engines and more specifically to cooling circuits between compressors and turbine blades.
Gas turbine engines operate by passing a volume of high energy gases through a plurality of stages of vanes and blades, each having an airfoil, in order to drive turbines to produce rotational shaft power. The shaft power is used to drive a compressor to provide compressed air to a combustion process to generate the high energy gases. Additionally, the shaft power is used to drive a generator for producing electricity, or to drive a fan for producing high momentum gases for producing thrust. In order to produce gases having sufficient energy to drive the compressor, generator and fan, it is necessary to combust the fuel at elevated temperatures and to compress the air to elevated pressures, which also increases its temperature. Thus, the vanes and blades are subjected to extremely high temperatures, often times exceeding the melting point of the alloys comprising the airfoils. High pressure turbine blades are subject to particularly high temperatures.
In order to maintain gas turbine engine turbine blades at temperatures below their melting point, it is necessary to, among other things, cool the blades with a supply of relatively cooler air, typically bled from the high pressure compressor. The cooling air is directed into the blade to provide impingement and film cooling. For example, cooling air is passed into interior cooling channels of the airfoil to remove heat from the alloy, and subsequently discharged through cooling holes to pass over the outer surface of the airfoil to prevent the hot gases from contacting the vane or blade directly. Various cooling air channels and hole patterns have been developed to ensure sufficient cooling of various portions of the turbine blade.
A typical turbine blade is connected at its inner diameter ends to a rotor, which is connected to a shaft that rotates within the engine as the blades interact with the gas flow. The rotor typically comprises a disk having a plurality of axial retention slots that receive mating root portions of the blades to prevent radial dislodgment. The siphoned compressor bleed air is typically routed from the compressor to the turbine blade retention slots for routing into the interior cooling channels of the airfoil. As such, the bleed air must pass through rotating and non-rotating components between the high pressure compressor and high pressure turbine. For example, cooling air is often drawn from the radial outer ends of the high pressure compressor vanes and routed radially inward through a support strut to the high pressure shaft before being directed radially outward for flow across the turbine rotor and into the turbine blade roots. Routing of the cooling air in such a manner incurs aerodynamic losses that reduce the cooling effectiveness of the air and overall gas turbine engine efficiency. Additionally, the bleed air must also pass through high pressure zones within the engine that exceed pressures needed to cool the turbine blades. There is, therefore, a continuing need to improve aerodynamic efficiencies in routing cooling fluid within cooling systems of gas turbine engines.